Turbine blade with multiple metering trailing edge cooling holes

• US 7,780,414 B1
• Filed: 01/17/2007
• Issued: 08/24/2010
• Est. Priority Date: 01/17/2007
• Status: Expired due to Fees

• Current Assignee: Florida Turbine Technologies, Inc. (Kratos Defense & Security Solutions Incorporated)
• Original Assignee: Florida Turbine Technologies, Inc. (Kratos Defense & Security Solutions Incorporated)
• Inventors: Liang, George
• Primary Examiner(s): Ward; Jessica L
• Assistant Examiner(s): Yoon; Kevin E
• Application Number:
  US11/654,126
• Time in Patent Office: 1,315 Days
• Field of Search: 164/369, 416/97.R, 415/115
• US Class Current: 416/97.00R
• CPC Class Codes: B22C 9/10 Cores; Manufacture or installation of cores breaker cores B22C9/084
• CPC Class Codes: F01D 5/187 Convection cooling
• CPC Class Codes: F05D 2230/21 by casting
• CPC Class Codes: F05D 2240/122 related to the trailing edge of a stator vane
• CPC Class Codes: F05D 2240/304 related to the trailing edge of a rotor blade
• CPC Class Codes: F05D 2250/185 serpentine-like
• CPC Class Codes: F05D 2260/201 by impingement of a fluid
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A ceramic core used to make a first or second stage turbine blade in an industrial gas turbine engine. The ceramic core includes a serpentine flow forming pierce with a trailing edge cooling supply channel forming piece on the trailing edge end. Three rows of metering holes are formed by a core piece in which a continuous flow channel piece on the tip portion and the root portion of the blade support the core piece that forms the metering holes and impingement cavity pieces such that the ceramic core is rigid and strong to prevent shear force and local bending of the core during casting will not break the core. A first and a second stage turbine blade is formed from the ceramic core and includes a forward flowing serpentine flow circuit for the first stage blade with the first channel of the serpentine flow circuit forming the trailing edge supply channel. A second stage blade includes an aft flowing serpentine flow circuit with the last channel forming the trailing edge supply channel. The three rows of metering holes allow for a gradual pressure drop from the high pressure trailing edge cooling supply channel and out the discharge holes or ducts along the edge of the blade.

Claims

1. A first stage or second stage turbine blade for an industrial gas turbine engine, the turbine blade comprising:

a leading edge and a trailing edge forming an airfoil extending from a platform to a blade tip;

a trailing edge region of the airfoil including a row of exit cooling holes extending from the platform to the blade tip;

the trailing edge region including three rows of metering holes that form a triple impingement cooling flow, the three rows of metering holes extending from the platform region to the blade tip;

a three-pass forward flowing serpentine flow cooling circuit with first leg being a cooling air supply channel located adjacent to trailing edge region;

the first leg of the serpentine flow circuit extending to an inner surface of the blade tip;

a lower continuous flow cooling air channel formed between the three rows of metering holes and the platform and connected to the first leg of the serpentine flow circuit;

an upper continuous flow cooling air channel formed between the three rows of metering holes and the inner surface of the blade tip and connected to the first leg of the serpentine flow circuit; and

,

the inner surface of the blade tip and the first row of metering holes and the lower and the upper continuous flow cooling air channels form a side of the first leg of the serpentine flow cooling circuit.

2. The first stage or second stage turbine blade of claim 1, and further comprising:

the blade tip includes tip cooling holes connected to the first leg of the serpentine flow cooling circuit and the upper continuous flow cooling air channel.

3. The first stage or second stage turbine blade of claim 1, and further comprising:

the blade tip forms a surface for a turn from the first leg to the second leg of the serpentine flow cooling circuit.

4. The first stage or second stage turbine blade of claim 1, and further comprising:

a leading edge cooling air supply channel connected to a leading edge impingement cavity through a row of metering and impingement holes;

the third leg of the serpentine flow cooling circuit being located adjacent to the leading edge cooling air supply channel; and

,

the third leg of the serpentine flow cooling circuit is not fluidly connected to the leading edge cooling air supply channel such that cooling air from one flows into the other.

5. The first stage or second stage turbine blade of claim 1, and further comprising:

the upper and lower continuous flow cooling air channels are of such size that a single ceramic core can be used to cast the first or second stage turbine blade of the industrial gas turbine engine with the trailing edge cooling circuit and the serpentine flow cooling circuit cast from the single ceramic core.

6. The first stage or second stage turbine blade of claim 1, and further comprising:

the upper and lower continuous flow cooling air channels form a bypass passage for the cooling air around the triple impingement metering holes.

7. The first stage or second stage turbine blade of claim 6, and further comprising:

the metering holes and the continuous flow channels are sized such that a cooling air pressure ratio across the entire trailing edge is around 1.6 and the cooling pressure ration across each row of metering holes is about 1.125.

8. A process for cooling a first stage or a second stage turbine blade of an industrial gas turbine engine, the process comprising the steps of:

supplying pressurized cooling air to a first channel located adjacent to a trailing edge region of the airfoil;

metering the pressurized cooling air from the first channel through three rows of metering holes that extend from near a platform to near a blade tip inner surface to provide triple impingement cooling;

passing some of the pressurized cooling air from the first channel around the metering holes through a lower continuous passage formed between the platform and the rows of metering holes to bypass the rows of metering holes;

passing some of the pressurized cooling air from the first channel around the metering holes through an upper continuous passage formed between the platform and the rows of metering holes to bypass the rows of metering holes;

passing the cooling air from the first channel through a second and third channel that forms a triple pass serpentine flow cooling circuit;

passing pressurized cooling air through a leading edge cooling air supply channel located adjacent to the third leg of the serpentine flow circuit; and

,

impinging the cooling air from the leading edge cooling air supply channel into a leading edge impingement cavity to provide near wall cooling to the leading edge of the airfoil.

9. The process for cooling a first stage or a second stage turbine blade of claim 8, and further comprising the step of:

discharging some of the cooling air from the first channel and the upper continuous passage through blade tip holes to provide cooling for the blade tip.

10. The process for cooling a first stage or a second stage turbine blade of claim 9, and further comprising the step of:

passing the cooling air from the first channel of the serpentine circuit that does not flow through the metering holes or the upper or lower continuous passages or the blade tip holes into the second channel of the serpentine circuit.